Means for supporting the nozzles of the combustion chambers of internal-combustion turbines



April 29, 1952 RUBBRA 2,594,808

. MEANS FOR SUPPORTING THE NOZZLES v0F THE COMBUSTION CHAMBERS 0FINTERNAL-COMBUSTION TURBINES Filed March 8, 1948 5 Sheets-Sheet l April29, 1952 A. A. RUBBRA 2,594,808 MEANS FOR SUPPORTING THE NOZZLES OF THECOMBUSTION INTERNAL-COMBUSTION TURBINES Filed March 8, l9

CHAMBERS OF 48 3 Sheets-Sheet 2 'll-l MM MJBBM 7% #77 April 29, 1952 r vA. A. RUBBRA 2,594,808 MEANS FOR SUPPORTING m: NOZZLES OF THE COMBUSTIONCHAMBERS OF INTERNAL-COMBUSTION TURBINES Filed March 8, 1948 3Sheets-Sheet 3 of the pressed air-to-a plurality: of combustion chambersvrotor 101' rotors and the to the intermediate Patented Apr. 29, 1952PATENT OFFICE 2,594,808 'Moms'roascrrommo-Tun NozzL-cs on 'THEU'QMBUS'IYION CHADIBERS OF IN TER- WhL-ooMBUsTIoN rURBINEs-ArthucAlexanderIRubbra,

v Littleover, Derby, England, 'ass'ignor to Rolls-R oyce Limited, Derby,

IEnglan'fl,ja'British company Application March 8, 1948;Sei-ialNoJ13Q689 :sImGrea-t Britain "comers. (o1. cams invention relates tolgas-iturbine-engines kin n which .acompressor delivers comand"the-productsof Tcombustion flow from the combustion chambers through a.ring of nozzle ;guide"vanes or turbine .stator-blading to one or moreturbine rotors which drive the compressor :and which the :mainstructure:for supporting jthe turbine rotor assembly comprises the eompressor"casing-and an iin'termediate casing lying within the ring of combustionohanibersand enclosing the drive shaft between the turbine compressorrotor, and also-in which the turbine stator assembly, including shroudring or ring ,"stator blading or guide vanes is'suppcrted fromtheintermediate casing.

HithertoTit has been the vpractice tolprovide nozzle-boxes .betweenthecombustion chambers and the z-inlet to the turbine, the nozzle-boxesbeing'circular at their forward or .inlet endswhere they receive thecombustion products from he' combustion equipment-land.substantiallyrec-V itangulariat their rearor outlet ends where they direct the}combustion, products on to the nozzle guide van assembly; theoutlet-ends-of the nozz'les were iarranged side -by-side to form acomplete annular ring corresponding to theannulus' :of'the nozzle guidevane assembly.

Inrsuchjrknown constructions the nozzles were welded to'gether'alongradially abutting cages at their outlet ends and further the inner andouter diametral edgesat the outlet ends were welded to-inner and outersupport rings. With "such a construction, it has been the practice to"support the inner ring from the intermediate casing referred to, whilstthe turbine shroud ring or rings (together with "thejet pipe) were-supported from the outer ring. As a consequence the weight of the jetpipe was taken through the welded nozzles and also, where the nozzleguide vanes were rigidly secured by welding or otherwise to their innerand outer shroud rings, through itheguide vanes.

In the specification of'U. SsPatent No.2,494;82'1 granted January 17,1950, there is described a construction in which the nozzle-boxes andnozzle guide-vanes are relieved of structural loading and which permitsof these components being so supported that they can expand freelywithin the assembly. This construction comprises a ring having a: seriesof apertures each arranged to receive and support the outlet end oi acombustion chamber, the apertured ring being secured I casing and to theouter ring of the nozzle-guide-vane assembly so that strucsolely throughthe ringQiThe ring conv an apertured ring for receiving and supp turalloads are tra-nsmi ed from one tothe provides a wall of ahQusing'whichisuppor .Tth'e nozzle-boxes. This apertured ring 'hasb'eeni'firoduced bya casting operation.

This invention seeks to provide an improved construction of such anaperturedring which is relatively simple to manufacture and which, whileretaining the desirable features of a'castccnstruction and apredetermined over-answer the outlet ends of the combustionchambefsiof agas-turbine engine of the hind desbrib' f the form of a castingWhich'h'as beeiij achine'd substantially over its wholesurface scia'stahave maximum thickness between the apertures and to taper inthicknessito'vvards its edges.

In one constructionof apierturedwring according to this invention, onesurface of the ring is machined to the form ofa part of a sphere havingits centre on the axis of the ring 'rein'ote from the plane thereof andthe othersurface ofj'th'e ring is machined by a turning operation sothat the ring has a substantially triangular crosssection between theapertures. .Reces'se's in be formed in the ring between the apertures sothat the ring has a web-like formation between the apertures.

By the invention, the weight of the ring may be kept to a minimum-havingregard to aprodetermined overall strength thereon n According to anotherfeature of this nvention, socket members are secured to the ring engagein the apertures to guide and support the outlet ends of the combustionchambers,- and, where the ring is formed with a part-spherical surface,-the socket members are preferably each formed with a flange having aspherical surface to co-operate with thespherical surfaceof the ring. 7

The invention is particularly applicable to the type ofgas-'turbine-engine in which: the com-lous tion chambers are disposed ina ring around the engine with their axes lyingin the surface or a conethe apex of which is'beyond the outlet ends of the combustion chambers,and in thiscas the "rearwardly by the the direction of flight Referringnow e bolted to the inner ring 20 nular outlet to the end to the tubularmember .23. applied to'the turbine shroud ring 22 by the exh'austassembly 3, for example due to its weight,

centre of the spherical surface of the ring will be the apex of thecone.

There will now be described a gas-turbinecngine of which the nozzle-boxassembly comprises one construction of apertured ring according to thisinvention, the description making reference to the accompanyingdrawings, in

view of the asingle-stage turbine generally indicated at H. a

A plurality of combustion chambers l2 receive air from the compressor H)by ducts Na and liquid fuel is burnt in these chambers. The products ofcombustion are delivered to the turbine H and then pass to atmosphere byway of an exhaust assembly l3. The exhaust gases are directed exhaustassembly relative to of the aeroplane for propulsion purposes.

to Figure 2, it will be seen that theturbine shaft Hi supports theturbine disc it? which carries blades IS. The shaft M is supported by abearing H carried by a structure 13 extending from an intermediatecasing 29 which is bolted to the compressor casing Hi. In this way theturbine rotor assembly is supported from the compressor casing It, theintermediate cas ing [9 and the structure 13 constituting a backbonestructure for the engine. a

The exhaust assembly I3 is bolted at 2! to a turbine shroud ring 22which in turn is bolted at 23 to an outer ring 26 between which and aninner ring 20 extend the fixed guide vanes 25 of the turbine.

' A ring 26, forming one part of a labyrinth seal. which is itselfbolted at 21 to a tubular member 28.

A nozzle-box assembly is located between the ou tlet ends of thecombustion chambers i2 and the guide vanes -25. The nozzle-box assemblycomprises a ring-like casting 29 having an apertured forwardly-facingwall 2% by which the outlet ends of the combustion chambers l2 are supported and a series of nozzle-boxes 3b which are supported at theirinlet ends by the casting 29 to receive the combustion gases from thecombustion chambers and which are supported at their outlet ends inbetween a flanged ring SI and the tubular member 28. It will beunderstood that the nozzle -boxes are circular at their inlet ends andare substantially rectangular at their outlet ends "to abut'laterally toform a substantially anguide vanes 25. The casting 29' is formed with anouter periphral wall 2% which extends rearwardly and um wardly to bebolted at 32 to the flange 32a on.

the ring 3! and the outer ring 14 of the guide vane assembly. and withan axial, forwardly-di rected flange 290 at its inner periphery which isbolted at 3-3 at its front end to the intermediatecasing 1'9" andstructure 18 and at 34 at itsrear In this way loads 23 and inner ringguide-vane assembly to the casting 29 which transmits them to theintermediate casing [9. These parts constitute the main load carryingstructure of the engine and the tubular member 20 which support theguide blades 25 are substantially unstressed. The guide are transmittedthrough the outer ring 2 of the entially by the outer rin combustionchamber the apertures, the

blades 25 are arranged to be located circumfer- 24 but to have radialfreedom with respect to the ring so that any loads on the rings 29, 2dare not imparted to the blades.

The nozzle-boxes 30 are provided at their inlet and outlet ends withspringy flanges 35 by which they engage with the casting 29 and theparts 3| and 28 so that the nozzle-boxes are also substantially freefrom loads in the main engine structure.

The combustion chambers l2, which are disposed in a ring around theshaft I4, have their inlet ends at a greater radius than their outletends and the axes of the combustion chambers lie in the surfaceof a conehaving its'apex 36 on the shaft axis and the apertured wall 29a of thecasting 25 is arranged to be substantially normal to the axes of thecombustion'chambers.

Referring now to Figures3 to 6, which'illustrate the casting 29 indetail, it will be; seen that the casting 29 is formed with a series ofcircular apertures 31 which are machined to have a cylindrical peripheryand therefore to' lie'substantially in a plane which is normal to-theaxis of the combustion chamber associated with it.

F The wall 29a is machined on its inner surface 38 to be part of asphere having its centre at the apex 3 6 of the cone in the surface ofwhich the axes lie.

The outer surface 3'3 of the wall 29a isf also machined by a simpleturning operation about the axis oi the ring 29 so that the wall 29athereof has .a thickness predetermined in' relation to thepart-spherical surface 38. It is arranged that the minimum thickness isobtained in the areas lying beyond the apertures 31, that is in theperipheral areas oft-he wall 2% and that the thickness is increased inthe region between the apertures in order to increase the strength ofthe ring where it is weakened by the presence of the apertures $1. Aswill be seen more clearly from Figure 5, the wall 25a has its maximumthickness between the apertures 3i and tapers smoothly towards itsperipheral edges so as to have a substantially triangular radialcross-section between the apertures.

In order to avoid giving the ring an excessive weight due to theincrease in thickness between ring 29 is preferably cast with recessesit) in the thickened portion so that the metal between each pair-ofapertures is in effect constituted by webs (see Figure 6) bounding theapertures circumferentially." It will be appreciated that this web-likeformation gives the 'ring 1 and 43 forming guides and sealing surfacesrespectively for the outlet ends of the combustion chambers and theinlet ends of the nozzle boxes. Each socket-member M is formed with aperipheral flange 44 having one surface of spherical form to-Jco-operatewith the surface 38' of the wall 29c; and the socket-members Allare-secured in position by studs .5 extending through the wall' 29a intothe flanges M. Flats 86 are machined on the front surface 39 of the wallto provide a seating for-the heads of the studs 45.

, receive and support the The outer peripheral Wall 2% and the innerperipheral flange 290 may also be machined to a desired thickness by asimple turning operation about the ring axis.

I claim:

1. An axial-flow gas-turbine comprising a turbine rotor; a stator casingenclosing the rotor; a

driving shaft on said turbine rotor; a plurality of combustion chambersdisposed in a ring around the said shaft with their axes intersectingthe turbine axis all at a single point; an intermediate casingsurrounding the shaft within the ring of combustion chambers; aring-like casing member connected to the intermediate casing and to thestator casing to transmit structural loads therebetween, said ring-likecasing member having one machined surface formed as a portion of asphere whose centre liesat said point, and having formed therein aplurality of apertures, one aperture for each combustion chamber spacedcircumferentially to correspond with the disposition of the outlet endsof the chambers, the angular position of said ring-like casing memberbeing such that the axes of the combustion chambers intersect at saidpoint, said ring-like casing member having a second machined surfaceopposite to said first machined surface, said second surface being ofsubstantially non-uniform convex contour so that the said two surfacescooperate to provide the greater thickness of the ring-like casingmember in the region lying between said apertures and the less thicknessin the peripheral regions of the ring-like casing member, and aplurality of apertured socket members one in each of said apertures tooutlet ends of the combustion chambers, said socket members each beingsecured to said casing member on the spherical side thereof by flangemeans formed on the socket member.

2. An axial-flow gas turbine as claimed in claim 1 wherein the socketmembers are formed with a flange having a surface machined to be part ofa sphere to co-operate with the spherical surface of the ring.

3. An axial-flow claim 1 wherein the gas turbine as claimed incross-section of the portion of the ring-like casing member lyingbetween the apertures, cone whose axis is the axis of the turbine, is ofU-shaped form.

4. An axial-flow gas-turbine comprising a turbine rotor; a stator casingenclosing the rotor; a driving shaft on said turbine rotor; a pluralityof combustion chambers disposed in a ring around the said shaft; anintermediate casing surrounding the shaft within the ring of combustionchambers; a ring-like casing member connected to the intermediate casingand to the stator casing to transmit structural loads therebetween, saidring-like casing member having one machined surface formed as a portionof a sphere whose centre lies on the axis of the turbine, and havingformed therein a plurality of circumferentially spaced apertures, oneaperture 'for each combustion chamber, said ring-like casing member faceopposite to said first machined surface, said second surface being ofsubstantially non-uniform convex contour and cooperating with said firstmachined surface to define areas of greater thickness of the ring-likecasing member between said apertures and of relatively less thickness inthe peripheral areas of the ring-like casing memher, and. a plurality ofsocket members one in each of said apertures to receive and support theoutlet ends of the combustion chambers, said ARTHUR ALEXANDER RUBBRA.

REFERENCES CITED The following references are of record in the file ofthis patent:

UNITED STATES PATENTS on a plane forming the surface of a having asecond machined sur-

